1. Field of the Invention
The present invention relates to a seal structure for a gas turbine, and more particularly to a seal separating surface connecting structure which reduces the amount of air leaking from a seal separating surface and improves sealing performance, and a structure which employs a brush seal so as to prevent a support plate of the brush seal from seriously contacting with a rotation side and which provides a small sealing interval so as to improve the sealing performance.
2. Description of Related Art
FIG. 9 is a cross sectional view which shows a general seal structure for a conventional gas turbine. In the drawing, reference numeral 61 denotes a stator blade, reference numeral 62 denotes an outer shroud thereof, and the outer shroud 62 is supported by a blade ring 70. Reference numeral 63 denotes an inner shroud, reference numeral 64 denotes a flange portion thereof and reference numeral 65 denotes a seal ring holding ring. The seal ring holding ring 65 is held by the flange portion 64 of the inner shroud 63, and supports a seal ring 66 in an inner side (a rotor side) thereof. Reference numerals 67 and 68 denote seal portions, and seal fins 67a and 67b are provided in a side of a rotor disc 69, thereby constituting a seal portion in opposition to the seal portion of the seal ring 66. Reference numeral 81 denotes a rotor blade, and reference numeral 82 denotes a platform thereof. The platform 82 is mounted on the rotor disc adjacent to the stator blade 61 and rotates together with the rotor. Reference numeral 71 denotes a tube for sealing air, which is provided within the stator blade 61, extends through the inner shroud 63 from the outer shroud 62 and is structured so as to introduce the sealing air within a cavity 75. Reference numerals 72, 73 and 74 denote spaces formed with respect to the adjacent rotor blade.
In the stator blade having the structure mentioned above, air 40, introduced from an inner portion of the sealing air tube 71 within the stator blade 61 to an inner portion of the cavity 75, passes through a hole 65a in the seal ring holding ring 65 and flows into the space 72 as shown by reference symbol 40a, and a part thereof flows out from the space 72 to the space 73 as leaking air from the seal portion 68 as shown by reference symbol 40b, passes between the platforms of the rotor blade adjacent to the inner shroud 13 as shown by reference symbol 40c and flows out to a main current gas passage.
Further, the air also flows out to the space 74 from the space 72 through the portion between the seal ring 66 and the seal portions 67a and 67b as shown by reference symbol 40e, and flows out from there through the portion between the inner shroud 63 and the platform 82 of the adjacent rotary blade 81 as shown by reference symbol 40f. A pressure within the cavity 75 and the spaces 72, 73 and 74 is increased in comparison with the outer main current gas passage due to the air flows mentioned above, thereby preventing the high temperature combustion gas from entering the inner portion. Accordingly, in order to increase sealing performance, it is necessary to increase the sealing performance of a labyrinth seal formed by the seal ring 66, the seal portion 67 on the rotor side and the seal portion 68 so as to reduce the amount of leaking air represented by reference symbols 40b, 40c, 40e and 40f. When it is possible to reduce the flowing-out air amount, it is possible to reduce the sealing air amount, so that it will be possible to improve the performance of the gas turbine as a whole.
FIG. 10(a) is a schematic view as seen from a line Zxe2x80x94Z in FIG. 9, which illustrates an upper half of the seal ring 66. Further, FIG. 10(b) is a schematic view taken along a line Wxe2x80x94W in FIG. 10(a). The seal ring 66 is formed in a circular ring shape and is separated into a plurality of portions, and in the drawing, the upper half portion is separated into three portions comprising portions (A), (B) and (C) (totally separated into four portions). Fins 66a, 66b and 66c are respectively mounted on the separated pieces (A), (B) and (C), which are opposed to the seal portions 67a and 67b in the side of the rotor disc 69. In the separated structure of the seal ring 66 mentioned above, a gap 51 is disposed between the separated pieces (A) and (B), a gap 52 is disposed between the separated pieces (B) and (C), and air flows out from an upstream space 72 (see FIG. 9) to a downstream space 74 (see FIG. 9) as the leaking air 60 flows through gap 51 as shown in FIG. 10B, so that the seal performance of the seal ring 66 is lowered.
As mentioned above, the seal structure of the conventional gas turbine is structured such that the seal ring 66 is separated into a plurality of portions, as mentioned above, and has a gap at a connecting end surface between the separated pieces of each of the seal rings. It is necessary to structure the gap such that a certain degree of gap is kept in connection with the thermal expansion of the seal ring. The gap is formed in a straight shape from the upstream side of the main current gas to the downstream side, and the sealing air leaks from the upstream side to the downstream side, so that the performance of the labyrinth seal constituted by the seal portion of the seal ring 66 and the seal portion 67 of the rotor disc 69 is reduced, and the amount of sealing air is increased, thereby providing an effect on the performance of the whole of the gas turbine.
Further, in recent years, in order to improve the seal performance of the stator side and the rotor side in the gas turbine for a flying machine or an industrial machine, the brush seal is going to be used. FIG. 11 is a representative cross sectional view of the industrial gas turbine, in which reference numeral 101 denotes a rotor blade, reference numeral 102 denotes a platform and reference numeral 103 denotes a seal pin within the platform 102, which is constituted by portions 103a, 103b and 103c. Reference numerals 102a and 102b denote a seal portion at both end portions in front and rear with respect to the axial direction of the platform 102. Reference numeral 104 denotes a shank portion and seal plates 106 and 107 are provided at the front and at the rear of the shank portion. Reference numeral 105 denotes a disc for a rotor, to which parts of each of the rotor blades 101 to 104 are mounted.
Reference numeral 91 denotes a stator blade, reference numeral 92 denotes an inner shroud and reference numeral 93 denotes an outer shroud. Reference numerals 92a and 92b denote the front and rear end portions with respect to an axial direction of the inner shroud 92. Reference numeral 94 denotes a cavity formed within the inner shroud 92, into which the sealing air is flowed through from the inner portion of the stator blade 91. Reference numeral 95 denotes a seal box, which holds a labyrinth seal 100 at one portion and a brush seal 101 at the other portion. Reference numerals 96 and 97 denote honeycomb seals provided at both end portions 92a and 92b of the inner shroud 92, respectively. Reference numerals 98 and 99 denote spaces formed with respect to each of the adjacent front and rear rotor blades, which correspond to flow passages for the sealing air.
In the gas turbine having the structure mentioned above, the sealing air is introduced into the cavity 94 from a sealing air passage (not shown) passing through the inner shroud 92 after passing from the side of the outer shroud 93 in the stator blade 91 through the inner portion of the blade. The air flows out into the space 99 from a hole (not shown) provided in the seal box 95. The air then passes through the honeycomb seal 96 provided at the end portion 92a of a the inner shroud 92, and flows out into the combustion gas passage. Further, a part of the sealing air flowing out from the hole in the seal box 95 also flows out to the forward space 98 through the brush seal 101 and the labyrinth seal 100 with respect to the disc 105. The air then flows out into the combustion gas passage through the honeycomb seal 97 provided at the front end portion 92b of the inner shroud 92.
As mentioned above, the sealing air is introduced within the cavity 94 through the inner portion of the stator blade 91, and the sealing air is introduced into the space 99, which is sealed by the honeycomb seal 96, from the cavity 94 and further into the forward space 98, sealed by the honeycomb seal 96, through the brush seal 101 and the labyrinth seal 100 with respect to the combustion gas passage, so that a pressure in the cavity 94 and the spaces 98 and 99 is increased higher than the pressure in the outer combustion gas passage, thereby preventing the high-temperature combustion gas from entering these areas.
FIG. 7 is an enlarged cross sectional view of the brush seal 101 corresponding to the seal portion of the gas turbine mentioned above. In the drawing, the brush seal 101 is mounted on the upstream side of the labyrinth seal 100 in the seal box 95 so as to constitute the seal portion with respect to the rotor disc 105. The brush seal 101 comprises a front portion support plate 31 and a rear portion support plate 32, a brush 33 comprising a multiplicity of narrow wires is fixed to a center portion thereof by a welding portion 35, and a front end thereof constitutes a seal portion in close contact with the disc 105. A gap 34 is provided between the front portion support plate 31 and the brush 33, and the structure is formed so as to move against a pressure in the upstream side. Widths Wf and Wa of support plate front ends 31a and 32a of the front portion and rear portion support plates 31 and 32 are set at about 2 mm, and the brush 33 is held at the front end portions.
FIG. 8 is a plan view of the brush seal 101, which is mounted after being separated into six portions in a circular ring shape, and an end surface thereof is brought into contact with the seal brushes adjacent to each other with an incline of 45 degrees in a rotational direction R with respect to a circular ring-like center line. The brush 33 is constituted by bundling a multiplicity of wires each having a diameter of 0.1 mm, and structured such that a width is set to about 2 mm and 1500 to 2000 wires are provided at each 25 mm length of the brush seal. The brush seal 101, structured in the above manner, is separated into six portions in a circumferential direction, and the front end of the brush 33 is in close contact with the disc 5 at the rotor side, thereby constituting the seal portion.
As mentioned above, the brush seal is used as the seal portion between the stationary portion and the rotary portion of the gas turbine. However; as shown in FIG. 7, the widths Wf and Wa of the front ends 31a and 32a of the support plate of the brush seal 101 are large at about 2 mm since they support the front end portion of the brush 33, so that the front end of the brush 33 deforms due to a thermal deformation of the disc 5 in the rotor side. In the case where the front ends 31a and 32a are brought into contact with the disc 105, the front ends 31a and 32a can cause serious damage on the surface of the disc 105 due to the significant rigidity thereof. Accordingly, the damage to the rotor seriously troubles the operation of the gas turbine, so that maintenance and replacement are performed on a large scale, and it is necessary to avoid the damage to the rotor as much as possible.
The present invention is made for the purpose of providing a seal structure for a gas turbine which designs a shape of a separation end portion of a seal ring, reduces the amount of air leaking from a separation portion, and improves a seal structure in the seal portion so as to improve sealing performance.
Further, the present invention is made for the purpose of improving the shape of a front end portion of a brush seal of a gas turbine, thereby avoiding double contact even when a rotor disc and a front end of the brush seal are brought into contact with each other so as not to seriously damage the rotor.
In order to achieve the objects mentioned above, the invention provides the following (1) to (5) items.
(1) A seal structure for a gas turbine comprising a seal ring holding ring supported on an inner shroud in a stationary blade, and a circular ring-like seal portion supported on the seal ring holding ring and surrounding a periphery of a rotor in a state of maintaining a predetermined gap obtained by taking thermal expansion into consideration between end portions of a plurality of separation pieces or contacting the end portions, thereby constituting a seal portion with the rotor. The circular ring-like seal portion is constituted by a brush seal at an upstream side of a main current gas and a labyrinth seal at a downstream side, and the end surfaces of each of the separation pieces of the brush seal and the labyrinth seal is formed in such a manner as to be in contact with each other at a cutting surface formed in a polygonal line.
(2) A seal structure for a gas turbine, as recited in the invention described in item (1), wherein the seal ring holding ring is separated into two portions at the upstream side and the downstream side, and one of them supports the brush seal and the other of them supports the labyrinth seal, respectively.
(3) A seal structure having a brush seal for a gas turbine comprising front and rear support plates mounted to a seal support portion inside a stationary blade of a gas turbine, a brush supported between front and rear support plates and a front end of the brush sealing with respect to a rotary portion in a rotor side. The front end portions of the front and rear support plates for supporting the front end portion of the brush have an axial thickness of 0.2 to 0.8 mm and are formed in a fin shape.
(4) A seal structure having a brush seal for a gas turbine as recited in item (3), wherein the front end portion of the support plate forms a fin shape by a smooth curved surface.
(5) A seal structure having a brush seal for a gas turbine as recited in item (3), wherein the front end portion of the support plate forms a fin shape by a straight tapered shape.
In the seal structure constructed in accordance with item (1) of the invention, since the upstream side of the circular ring-like seal portion is constituted by the brush seal, sealing performance is improved in comparison with the conventional labyrinth seal. Further, since the end surface of each of the separation pieces in the circular ring-like seal portion has the shape contact with each other at the cutting surface formed in the polygonal line, air that is going to flow out to the downstream side from the upstream side through the gap is hard to be leaked in comparison with the conventional straight flow passage because the flow passage becomes a bypass passage and a flow resistance is increased, even when a gap is generated due to a thermal deformation. Accordingly, the amount of air amount is leakage is reduced, sealing performance is improved, and the amount of air is reduced, so that performance of the whole of the gas turbine is improved.
In item (2) in accordance with the is invention, the seal ring holding ring is separated into two portions and one of the portions supports the brush seal and the other supports the labyrinth seal, respectively. Accordingly, in addition to the effect of item (1) mentioned above, there is an advantage in that an assembling and working performance thereof can be improved since the assembly of each of the brush seal and the labyrinth seal can be separately performed.
In the brush seal constructed in accordance with item (3) of the invention, the front end portions of the front and rear support plates for supporting the front end portion of the brush have a thickness of 0.2 to 0.8 mm which is thinner by 2 mm with respect to the conventional case, and is formed in a fin shape. In accordance with the structure mentioned above, the stationary side and the rotor side are thermally deformed, so that even when the front end of the support plate is brought into contact with the rotor side, the front end of the support plate deforms due to the thin fin shape and elasticity. Accordingly, the support plate side is deformed without being in double contact with the rigid body in the conventional art, or the support plate itself is broken, thereby avoiding damage to the rotor side. Further, since the double contact can be avoided and an elastic force can be added, the seal interval can be made smaller than in the conventional one, so that sealing performance can be improved.
In item (4) in accordance with the invention, since the front end portion of the support plate forms the fin shape by the smooth curved surface, and in item (5), the front end portion of the support plate forms the fin shape by the straight tapered shape, respectively, the front end portion gradually changes so as to form a narrow front end, so that the strength of the front end portion can be secured.